The performance of a gas turbine engine cycle, whether measured in terms of efficiency or specific output is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbine at the highest possible temperature. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature will always produce more specific thrust (e.g. engine thrust per unit of air mass flow). However, as turbine entry temperatures increase, the life of an uncooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high pressure (HP) turbine gas temperatures are now much hotter than the melting point of the blade materials used and in some engine designs the intermediate pressure (IP) and low pressure (LP) turbines are also cooled. During its passage through the turbine the mean temperature of the gas stream decreases as power is extracted. Therefore the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the HP stage(s) through the IP and LP stages towards the exit nozzle.
Internal convection and external films are the prime methods of cooling the aerofoils. HP turbine nozzle guide vanes (NGVs) consume the greatest amount of cooling air on high temperature engines. HP blades typically use about half of the NGV flow. The IP and LP stages downstream of the HP turbine use progressively less cooling air.
Blades and vanes are cooled by using high pressure (HP) air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 700 and 1000 K. Gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Extracting coolant flow therefore has an adverse effect on engine operating efficiency. It is thus important to use this cooling air as effectively as possible.
Previous cooling arrangements typically involve bleeding coolant from the tip sections of the aerofoil cooling scheme to cool the shroud. The cooling air passes radially into the core printout which has been blanked off at its extremity using a welded plug that is drilled with a smaller hole to allow the passage of dust and dirt and prevent contamination and/or blockage of the shroud cooling holes. An array of small diameter cooling holes are drilled into the edge of the shroud connecting with the cavity created by blanking the core printout. The spent aerofoil cooling flow passes radially through the core printout (typically referred to as the ‘Chimney Pot’) and then circumferentially through the array of cooling holes. Finally, the coolant emerges into a cavity machined into the pressure surface edge of the shroud and impinges onto the suction surface edge of the neighbouring shroud. The emerging coolant mixes with the hot gas that leaks over the labyrinth seal fins located on the upper surface of the shroud. This leaking hot gas and cooling air mixture is highly swirled as a consequence of being trapped between the rotating blade upper surface and the stationary shroud seal segment.
FIG. 1 illustrates a typical prior cooling arrangement between adjacent shrouded blades a gas turbine engine. Thus, as described above, coolant air flows in the direction of arrowheads 2 through the core printout and drilled dust hole 3 in order to cool the shroud's upper surface and fins 4. In accordance with this prior arrangement, as indicated above, a number of discrete passages 5 are formed below shroud surface 6 in order that a proportion of the coolant flow passes along the passage 5 to be output as an ejected coolant flow 17. An end of the shroud surface 6 is generally cut back to create an edge 18 in order to facilitate coolant flow 17.
As indicated above generally there is a cascade in necessary cooling between turbine stages in a gas turbine engine.
In the arrangement 1 depicted in FIG. 1 adjacent shrouds 6, 7 are arranged such that the coolant flow 17 cools a suction surface 8 of the shroud 7 adjacent to the shroud incorporating the shroud surface 6. It will be appreciated that a further coolant flow 12 is provided within the adjacent blade shroud in order that a similar cooling regime in accordance with the previously described arrangement is achieved along a passage 5. In such circumstances, it will be appreciated that adjacent blade shroud is arranged to provide cooling in a cascade. Furthermore, a number of passages 5 will be provided around the circumference of the assembly in order to provide adequate cooling of the shroud surfaces and in particular the suction surface 8 of each shroud.
With the above described prior cooling arrangements, although the pressure wall surface 16 of the shroud 6 is effectively cooled by the embedded cooling passages 5, the suction wall surface 8 of the shroud 7 is only effectively cooled on an extreme edge 9 where the impingement coolant jets 17 strike, leaving the remainder of the shroud suction surface 8 inadequately cooled. The spent impingement coolant becomes entrained by the hot swirling, gas/coolant mixture and over tip leakage flow in the cavity above the surfaces.
In view of the above, it has been found in previous cooling arrangements that it is necessary to provide additional cooling of the suction surface 8. As indicated, cooling is a drain upon engine efficiency and therefore diminishes performance. In FIG. 1 in the above circumstances, it will be appreciated that the portion of the suction surface marked by stars 10 will generally be poorly cooled. In such circumstances, with hot gas 11 washing the inner surface of the shrouds 6, 7 it will be understood that there may be inadequate cooling of the suction surface 8 of the shroud 7 in the region of the stars 10.